It is generally well-known in the art of satellite control to provide an unsupported proof mass aboard a satellite, shielded from external non-gravitational forces so that it follows a purely gravitational orbit during operation of the satellite, and means responsive to motion of the satellite relative to the proof mass for controlling thrusters onboard the spacecraft which force the spacecraft to also follow the gravitational orbit, free from the effects of external surface forces such as solar radiation pressure and atmospheric drag. As a result, satellite position in orbit is predictable well in advance, thus significantly increasing the value of the satellite for navigational purposes for example.
One major problem which was encountered in the use of early disturbance compensation systems of this type involved in the need for very accurately determining mass attraction forces between satellite-carried components and the proof mass, in order to prevent such mass attraction forces from influencing the proof mass and thereby providing faulty operation of the system. In an effort to overcome the problem of mass attraction forces, U.S. Pat. No. 3,785,595 proposed the use of movable compensation masses or chargeable magnets (reacting with a diamagnetic proof Mass), on three orthogonal axes relative to the proof mass, which could be controlled by commands from the ground tracking station, once the spacecraft was in orbit, to set up forces which counterbalance the mass attraction forces.
U.S. Pat. No. 4,170,904 discloses a single-axis disturbance compensation system that is much simpler, for example, than the previous three-axis compensation system disclosed in U.S. Pat. No. 3,785,595. The single-axis compensation system of U.S. Pat. No. 4,170,904 utilizes a proof mass fabricated from electrically conductive material which is suspended electromagnetically, by eddy current forces, to counterbalance mass attraction forces due to the spacecraft, while moving with essentially no friction along an axis aligned with the velocity vector of the spacecraft. The proof mass is housed in an enclosure attached to the spacecraft, shielded from all non-gravitational exterior forces such as solar radiation pressure, atmospheric drag and electrostatic attraction forces, so that the proof mass follows a purely gravitational orbit. A controlled magnetic biasing field is generated within the enclosure adjacent to the proof mass and exerts a preselected and controlled eddy current biasing force level on the proof mass in order to balance any along axis constant component of proof mass disturbance force. Essentially, this biasing system comprises two oppositely wound coils each configured to provide a constant force over the proof mass range of motion. In order to assure that the proof mass always reacts in a constant fashion to the biasing magnetic field, a thermal control system is provided in order to assure constant resistivity of the proof mass materials.
When the satellite is placed into orbit and is controlled to assume its orbital configuration, the proof mass is positioned at the center of mass of the spacecraft in alignment with the velocity vector. An optical detection system then monitors any movement of the proof mass both along and transverse to the velocity vector, as results from external forces on the satellite due to solar radiation pressure and atmospheric drag. The output of the optical detection system is utilized to operate thruster control apparatus which force the satellite to maintain a substantially constant relative position with respect to the proof mass and whereby the satellite is also caused to follow the purely gravitational orbit.